When a vehicle is launched from a stationary launch platform and begins its launch trajectory, one of the most important considerations is to ensure that the vehicle does not rotate excessively about its lateral (pitch) axis. In other words, it is essential to stabilize the pitch of the vehicle. Thus, all launch vehicles are equipped with some type of pitch stabilization system. This is true for ejection seats as well as for other vehicles such as missiles and spacecraft. When an ejection seat is ejected from the aircraft cockpit by a catapult or other equivalent means, the windblast effect will push the top of the seat aft of the vehicle while the bottom of the seat is still constrained by the guide rails, producing a rotational moment on the seat and causing an undesirable aft pitch angle. To counter that tendency, the typical prior art pitch stabilizer consists of a gyroscope rotor which precesses in response to a change in pitch of the seat. This rotor is linked by some interconnection means with a stabilization thruster, which rotates in response to the precession of the gyro rotor. The change in thruster nozzle direction provides a counter-thrust to the pitch force and acts to stabilize the pitch of the seat. This system is very effective for its intended purpose.
The problem is that for the above system to work, the gyro rotor must be spun up to a speed sufficient for it to accurately sense pitch rate and precess accordingly, so that the thruster is rotated to counter the change in pitch. The problem of pitch change will be the most severe immediately after the cockpit breakaway point, so the gyro must be fully operational at that time. Since the rotor will be at rest until seat ejection occurs, the process of spinning up the rotor to operating speed must occur very rapidly, being accomplished prior to the time when the seat fully breaks away from the cockpit.
In prior art systems, the gyro rotor has been spun up using some type of on-board actuator. Most typically, this comprises an initiator charge, which is activated at the start of the ejection sequence. In one particular prior art system, the initiator is mounted on the seat and the gyro rotor is a toothed rotor, intermeshingly engaged with a toothed rack, also mounted on the seat. When the initiator is actuated, the rack is driven along the gyro rotor, the toothed interface acting to rotate the rotor. By the time the gyro rotor is disengaged from the end of the rack, it is at operating speed. One problem with a system such as this one is its weight. In an aircraft environment, each additional pound exacts a penalty in fuel cost and performance and the weight of the additional actuator is a consideration. Furthermore, it creates another complexity in the ejection sequence. Should the initiator fail, and the gyro rotor not be spun up to operating speed, the pitch stabilizer will be non-operational and the ensuing ejection could well be catastrophic to the seat occupant. What is needed is a simple, mechanically reliable, lightweight actuation system for ensuring that the gyro rotor is operational every time it is necessary to eject.